In order to increase the efficiency and the performance of gas turbine engines so as to provide increased thrust-to-weight ratios, lower emissions and improved specific fuel consumption, engine turbines are tasked to operate at higher temperatures. The higher temperatures reach and surpass the limits of the material comprising the components in the hot section of the engine and in particular the turbine section of the engine. Since existing materials cannot withstand the higher operating temperatures, new materials for use in high temperature environments need to be developed.
As the engine operating temperatures have increased, new methods of cooling the high temperature alloys comprising the combustors and the turbine airfoils have been developed. For example, ceramic thermal barrier coatings (TBCs) have been applied to the surfaces of components in the stream of the hot effluent gases of combustion to reduce the heat transfer rate, provide thermal protection to the underlying metal and allow the component to withstand higher temperatures. These improvements help to reduce the peak temperatures and thermal gradients of the components. Cooling holes have been also introduced to provide film cooling to improve thermal capability or protection. Simultaneously, ceramic matrix composites have been developed as substitutes for the high temperature alloys. The ceramic matrix composites (CMCs) in many cases provide an improved temperature and density advantage over metals, making them the material of choice when higher operating temperatures and/or reduced weight are desired.
A number of techniques have been used in the past to manufacture hot section turbine engine components, such as turbine airfoils using ceramic matrix composites. One method of manufacturing CMC components, set forth in U.S. Pat. Nos. 5,015,540; 5,330,854; and 5,336,350; incorporated herein by reference in their entirety and assigned to the assignee of the present invention, relates to the production of silicon carbide matrix composites containing fibrous material that is infiltrated with molten silicon, herein referred to as the Silcomp process. The fibers generally have diameters of about 140 micrometers or greater, which prevents intricate, complex shapes having features on the order of about 0.030 inches, such as turbine blade components for small gas turbine engines, to be manufactured by the Silcomp process.
Other techniques, such as the prepreg melt infiltration process have also been used. However, the smallest cured thicknesses with sufficient structural integrity for such components have been in the range of about 0.030 inch to about 0.036 inch, since they are manufactured with standard prepreg plies, which normally have an uncured thickness in the range of about 0.009 inch to about 0.011 inch. With standard matrix composition percentages in the final manufactured component, the use of such uncured thicknesses results in final cured thicknesses in the range of about 0.030 inch to about 0.036 inch for multilayer ply components, which is too thick for use in small turbine engines.
Complex CMC parts for turbine engine applications have been manufactured by laying up a plurality of plies. In areas in which there is a change in contour or change in thickness of the part, plies of different and smaller shapes are custom cut to fit in the area of the contour change or thickness change. These parts are laid up according to a complicated, carefully preplanned lay-up scheme to form a cured part. Not only is the design complex, the lay-up operations are also time-consuming and complex. Additionally, the areas of contour change and thickness change have to be carefully engineered based on ply orientation and resulting properties, since the mechanical properties in these areas will not be isotropic. Because the transitions between plies along contour boundaries are not smooth, these contours can be areas in which mechanical properties are not smoothly transitioned, which must be considered when designing the part and modeling the lay-up operations.
FIG. 1 depicts an exemplary uncoated airfoil (uncooled) 10. In this illustration the airfoil 10 comprises a ceramic matrix composite material. The airfoil 10 includes an airfoil portion 12 against which a flow of gas is directed. The airfoil 10 is mounted to a disk (not shown) by a dovetail 14 that extends downwardly from the airfoil portion 12 and engages a slot of complimentary geometry on the disk. The airfoil 10 does not include an integral platform. A separate platform can be provided to minimize the exposure of the dovetail 14 to the surrounding environment if desired. The airfoil has a leading edge section 18 and a trailing edge section 16. Such a composite airfoil is fabricated by laying up a plurality of plies.
FIG. 2 is a prior art illustration (perspective) of how such a composite airfoil of FIG. 1 has been laid up. FIG. 3 represents a front view of the lay-up of these pre-preg plies. The airfoil 10 comprises a plurality of pre-preg plies 40 arranged around a centerplane 24. There are a number of root (pre-preg) plies 41 and smaller (pre-preg) plies 42 arranged between larger (prepreg) plies 40, 44. Referring back to FIG. 1, the smaller plies, in particular root plies 41, are required to provide the dovetail geometry. In addition, each of the plies 40 includes tow that is oriented in a predetermined direction. Of course, care must be taken to not only provide the proper size ply in the proper location, but also to properly orient the tow direction of each of the plies.
Still other techniques attempt to reduce the thickness of the prepreg plies used to make up the multi-layer plies by reducing the thickness of the fiber tows. Theoretically, such processes could be successful in reducing the ply thickness. However, practically, such thin plies are difficult to handle during part manufacturing, even with automated equipment. This can result in wrinkling of the thin plies, a manufacturing defect that can result in voids in the article, and a deterioration of the mechanical properties of the article, and possible ply separation. In addition, problems arise, as airfoil hardware requires the ability to form small radii and relatively thin edges, features required in smaller articles, such as narrow chord airfoils. The high stiffness of the fibers, typically silicon carbide, in the prepreg tapes or plies, can lead to separation when attempting to form the plies around tight bends and corners with small radii. The fiber coatings may also crack or be damaged. This leads to a degradation in the mechanical properties of the article in these areas with resulting deterioration in durability.
What is needed is a method of manufacturing CMC turbine engine components that permits the manufacture of features having a thickness, particularly at the edges, in the range of about 0.015 inch to about 0.021 inch, as well as small radii, the radii also in the range of less than about 0.030 inches. In addition, a method of manufacturing CMC turbine engine components having features with a thickness less than about 0.021 inch is also needed.